U.S. patent number 5,281,084 [Application Number 07/552,281] was granted by the patent office on 1994-01-25 for curved film cooling holes for gas turbine engine vanes.
This patent grant is currently assigned to General Electric Company. Invention is credited to Mark E. Noe, Robert Proctor.
United States Patent |
5,281,084 |
Noe , et al. |
January 25, 1994 |
Curved film cooling holes for gas turbine engine vanes
Abstract
A method and apparatus for film cooling of an aerodynamically
shaped airfoil uses a plurality of curved slots extending through
the airfoil in an area upstream of the high curvature region of the
airfoil, i.e., in an area of low Mach number of the gas stream
passing over the airfoil surface. The curved slots are configured
to inject cooling air at an angle of about 16.5 degrees. The
cooling air is injected at a blowing ratio in excess of 1.0 and yet
is effective to form a film on the vane surface.
Inventors: |
Noe; Mark E. (Cincinnati,
OH), Proctor; Robert (West Chester, OH) |
Assignee: |
General Electric Company
(Cincinnati, OH)
|
Family
ID: |
24204673 |
Appl.
No.: |
07/552,281 |
Filed: |
July 13, 1990 |
Current U.S.
Class: |
415/115 |
Current CPC
Class: |
F01D
5/186 (20130101); F05D 2250/71 (20130101) |
Current International
Class: |
F01D
5/18 (20060101); F01D 009/04 () |
Field of
Search: |
;415/115,116
;416/96R,97R |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
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|
|
|
|
|
0079285 |
|
May 1983 |
|
EP |
|
0182302 |
|
Sep 1985 |
|
JP |
|
0032103 |
|
Feb 1988 |
|
JP |
|
2184492A |
|
Jun 1987 |
|
GB |
|
Other References
S Stephen Papell et al.; "Influence of Coolant Tube Curvature on
Film Cooling Effectiveness as Detected by Infrared Imagery"; NASA
Technical Paper No. 1546; 1979; pp. 3-16..
|
Primary Examiner: Look; Edward K.
Assistant Examiner: Verdier; Christopher
Attorney, Agent or Firm: Squillaro; Jerome C. Davidson;
James P.
Claims
What is claimed is:
1. A vane for a gas turbine engine, comprising:
an airfoil section having a generally convex suction surface
terminating in a trailing edge of the airfoil section, a generally
concave pressure surface opposite the suction surface and coupled
thereto at the trailing edge, the airfoil section further having a
relatively blunt leading edge coupling the suction surface to the
pressure surface through a transition region with a high curvature,
an enclosed chamber being defined by said leading edge, said
pressure surface and said suction surface within said airfoil
section;
a plurality of vent holes penetrating said airfoil section for
passing a cooling fluid from within said chamber to said surface of
said airfoil, said vent holes comprising arcuate passages extending
through said leading edge adjacent said high curvature transition
region for directing cooling fluid toward said suction surface at
an injection angle less than about 25 degrees, said cooling fluid
having a mass flow rate such that the blowing ratio at the vane
surface is greater than 1.0.
2. The vane as recited in claim 1 wherein the curved vent holes
have a radius of curvature of about 0.675 inches.
3. The vane as recited in claim 1 wherein the injection angle is
about 16.5 degrees.
4. The vane as recited in claim 1 wherein the blowing ratio is
about 1.2.
5. The vane as recited in claim 1 wherein a line tangent to one of
said vent holes and a line tangent to said trailing edge of said
vane form an angle therebetween of less than 90 degrees.
6. A method of cooling a vane in a gas turbine engine, the vane
having an airfoil section exposed to a stream of high temperature
combustion gases in the gas turbine engine and the airfoil section
including a relatively broad and blunt leading edge, a convex
shaped suction surface, a set of arcuate cooling air injection
holes in the leading edge of the airfoil section upstream of the
suction surface, and a chamber within the airfoil section
communicating with the cooling air injection holes, the method
comprising the steps of:
injecting cooling air from the chamber through the cooling air
injection holes onto the relatively broad and blunt leading edge
with an injection angle of less than about 25 degrees and
establishing a blowing ratio greater than about 1.0 in response to
the injecting step.
7. The method of claim 6 and further including the step of forming
the arcuate shaped air injection holes with a radius of curvature
of about 0.675 inches.
8. The method of claim 6 wherein the injecting step comprises
injecting cooling air at an injection angle of about 16.5
degrees.
9. The method of claim 6 wherein the step of adjustably
establishing a blowing ratio comprises the step of establishing a
blowing ratio of about 1.2.
Description
The present invention relates to vanes for gas turbine engines and,
more particularly, to vanes having hollow airfoil sections with
vent holes for cooling.
BACKGROUND OF THE INVENTION
The high temperature of inlet gas stream air entering high pressure
turbine nozzles and flowing over outer surfaces of individual vanes
of the nozzles in a gas turbine engine has required cooling of the
vane airfoil sections in order to maintain vane temperatures within
the present material capability. Cooling is commonly provided by
forming the vanes as hollow airfoils and providing vent holes from
the hollow interior through which a cooling gas, typically air, is
forced. The gas desirably forms a film over at least a portion of
the airfoil surface and thereby cools or at least insulates such
surface. The film cooling injection location is extremely important
on the suction side (convex surface) of the airfoil where the hot
gas stream can become supersonic. Performance considerations have
driven film cooling to be introduced on the airfoil surface at
locations where the hot gas stream has a low velocity and near the
leading edge of the airfoil section. The selection of cooling film
injection locations is a trade-off between performance and cooling
of the airfoil. Performance losses are directly proportional to the
square of the main stream Mach number at the injection locations.
Therefore, the impact on engine performance is significantly
different when comparing performance when coolant is injected in a
region where the Mach number is about 0.3 as opposed to injection
in a region where the Mach number is about 1.0. However, when
injection occurs in a low Mach number region, the cooling film may
degrade to a point of being ineffective prior to reaching the vane
trailing edge. In order to compensate for such degradation, it is
necessary to increase the flow of coolant, but such increased flow
adversely affects the temperature profile out of the combustor and
adversely affects engine performance. Accordingly, coolant
injection is often a trade-off of performance against cooling and
component life.
With some high curvature airfoil sections, the gas film or vent
holes are oriented angularly so as to reduce the gas film injection
angle. The reduced angle improves the ability of the film to flow
along the airfoil surface. If the film does not flow along the
surface, i.e., if it is dissipated in the gas stream, then cooling
is ineffective. Film blow-off occurs if the strength of the
injected coolant relative to the strength of the gas stream, i.e.,
the blowing rate, is incorrect for the coolant injection angle. It
has also been proposed to turn the cooling gas through a large
angle, e.g., between 135 and 165 degrees, using a curved admission
tube before injecting the cooling gas at an angle of between about
15 and 45 degrees with respect to the airfoil surface, to try to
force the film to remain on the vane surface over greater
distances. However, this arrangement has been applied to airfoils
having relatively continuously curved suction sides which do not
introduce rapid velocity changes. More particularly, this proposed
arrangement has been demonstrated to be effective only for blowing
rates of between about 0.37 and 0.70. For blowing rates above 0.70,
the curved tube was found to be less effective in film cooling than
straight tube injection. This above approach is discussed in detail
in NASA Technical Paper 1546 published in 1979 and entitled
"Influence of Coolant Tube Curvature in Film Cooling Effectiveness
as Detected by Infrared Imagery", by Papell, Graham, and Cageao. In
general, it is believed that blowing ratios greater than 1.1 are
less effective in film cooling.
The development of blunt leading edge airfoils creates more severe
film cooling requirements. With such airfoils, a high curvature
section exists immediately downstream of the normal film injection
point. Conventional injection processes are ineffective to maintain
the cooling film on the airfoil surface over such high curvature
regions. Furthermore, the velocity of the high temperature gases
over high curvature regions approaches supersonic velocities and
contributes to the degradation of the cooling film due to large
free stream turbulence.
SUMMARY OF THE INVENTION
It is an object of the present invention to provide a method and
apparatus for overcoming the above and other disadvantages
associated with film cooling of blunt airfoils in gas turbine
engines.
It is another object to provide a method and apparatus for cooling
of blunt airfoils which increases the effectiveness of film
cooling.
In one form of the invention, there is provided a vane for a gas
turbine engine nozzle which has an airfoil section with a broad,
blunt leading edge having a region of high curvature transitioning
from the leading edge to a convex shaped suction surface. A
plurality of vent holes are formed in the airfoil for conveying a
cooling gas from the hollow interior of the airfoil to the outer
surface thereof. At least some of the vent holes are located in the
broad leading edge of the airfoil immediately upstream of the high
curvature region such that cooling gas can be injected where the
velocity of the high temperature gas stream flowing along the vane
is relatively low. These vent holes are formed with an arcuate
shape through the airfoil wall so that the injection angle of the
cooling gas is less than 25 degrees and preferably about 16
degrees. The arcuate or curved vent holes serve to direct the
cooling gas downward along the airfoil surface and concurrently aid
in convection cooling of the airfoil by extending the length of the
holes through the airfoil wall. In addition, the blowing ratio can
be increased to values greater than 1.0 to obtain effective
cooling.
BRIEF DESCRIPTION OF THE DRAWINGS
For a better understanding of the present invention, reference may
be had to the following detailed description taken in conjunction
with the accompanying drawings in which:
FIG. 1 is a simplified partial cross-sectional view of an exemplary
gas turbine engine illustrating the location of the turbine vanes
to be cooled;
FIG. 2 is a simplified perspective view of a turbine vane of the
prior art;
FIG. 3 is a cross-sectional view taken through a turbine vane of
the type shown in FIG. 2; and
FIG. 4 is a cross-sectional view taken through a turbine vane
having a blunt leading edge and incorporating film cooling in
accordance with the present invention.
DETAILED DESCRIPTION OF THE INVENTION
FIG. 1 illustrates a triple spool front fan high-bypass ratio
ducted fan gas turbine engine 10 with which the present invention
may be used. The engine 10 includes a ducted fan 12, intermediate
and high pressure compressor sections 14 and 16, respectively, a
combustion chamber 18, a turbine stage 20, and an exhaust nozzle
22. The turbine stage 20 may be divided into high, low, and
intermediate sections for providing power to the fan 12 and
compressor sections 14, 16 through corresponding elements of a
central shaft 24. Shaft section 24A connects the final turbine
disks 20A to fan 12, shaft section 24B connects turbine disk 20B to
compressor section 14, and shaft section 24C connects turbine disk
20C to compressor section 16. Air compressed by fan 12 and the
compressor sections 14, 16 is mixed with fuel and combusted in
combustion chamber 18. The combustion products expand through the
turbine stage 20 and are exhausted through nozzle 22. Propulsive
thrust is provided by air moved outside the engine by the fan 12
coupled with some thrust provided by exhaust from the nozzle
22.
The turbine stage 20 includes a plurality of annular rows of
circumferentially spaced and radially extending nozzle guide vanes
26. Referring to FIG. 2, each vane 26 comprises an airfoil 28
having a radially inner platform 30 and a radially outer platform
32. The platforms 30 and 32 of adjacent vanes 26 cooperate with
each other as shown in FIG. 2 to define radially inner and outer
boundaries of a portion of the gas flow path through the turbine
stage 20. The airfoils 28 serve to direct the high temperature gas
stream from the combustion chamber 18 onto annular rows of rotor
blades coupled to respective sections of shaft 24. FIG. 3 is a
cross-sectional view taken through one of the airfoils 28 and
illustrates a prior art arrangement of cooling air holes 36 between
a hollow interior 34 and selected areas of the outer surface of the
airfoil. Cooling air delivered to the hollow interior 34 of the
airfoil and exhausted through the vent holes 36 flows along the
outer surface of the airfoil forming a film which cools the outer
surface and insulates it from the high temperature combustion
gases. The cooling air is generally supplied by tapping it from air
passing through the compressor section 16 in a manner well known in
the art.
The airfoil illustrated in cross-section in FIG. 3 represents a
typical prior art nozzle blade in which the airfoil has a
relatively continuous arc of curvature over its convex or suction
surface 38 extending from a relatively aerodynamic leading edge to
the trailing edge 42. The shape of the concave or pressure surface
44 is approximately the same as the suction surface 38. With such
smooth, continuously curved surfaces, it is relatively easy to
provide film cooling through use of substantially straight holes 36
passing through the walls 46. Some of these holes 36 may be
angularly oriented so that the cooling air is directed in the
direction of flow of the hot gas stream.
Film cooling is not primarily intended as protection of the surface
at the point of injection but rather as protection of the surface
at a region downstream of the injection location. The injection of
a cooling gas (air) into the boundary layer with film cooling may
be considered to produce an insulating layer or film between the
surface to be protected and the hot gas stream flowing over the
surface. The film layer also acts as a heat sink to lower the mean
temperature in the boundary layer adjacent the surface. As
described above, there is a trade-off between engine performance
and cooling air injection. If sufficient cooling air is not
injected onto the vane surface, the coolant will be dissipated too
quickly and will not be effective to protect the vane surface. If
the cooling air is injected at too high a rate, blow-off can occur.
This phenomenon occurs when the cooling flow drives away from the
vane surface because of its strength thus allowing the hot gas
stream to remain in contact with the surface, i.e., no insulation
layer is formed. Blowing ratio is a measure of the strength of the
injected cooling gas or air relative to the strength of the hot gas
stream. High blowing ratios are characteristic of blow-off. In
general, a blowing ratio in the order of 1.1 is characteristic of a
coolant injection rate which is ineffective, i.e., the coolant does
not form a surface film and degrades rapidly. Turbulence at the
surface of the airfoil due to abrupt shape (curvature) change also
contributes to such film degradation.
Studies have shown that improvement in film cooling can be somewhat
realized by increasing the flow of cooling air. However, it is
generally accepted that a blowing ratio (which compares the mass
flow per unit area of cooling air to the mass flow per unit area of
hot gases) cannot exceed about 1.0. The aforementioned NASA
Technical Paper 1546 compared the effectiveness of curved coolant
injection tubes to straight tubes and found that at blowing ratios
above 0.70, the effectiveness of curved coolant injection decreased
to a point where it became less effective than straight tube
injection. This, it is generally believed that film cooling is not
effective at blowing ratios above 1.0. More particularly, at
blowing ratios of about 1.1, the velocity of the cooling air is
sufficiently strong to detach itself from the surface and blow into
the hot gas stream.
Turning now to FIG. 4, there is shown a cross-sectional view of a
more recent design for a nozzle vane. The vane, indicated generally
at 48, has a broad, blunt leading edge 50, a convex shaped suction
surface 52, a concave shaped pressure surface 54, and a trailing
edge 56. While this vane airfoil configuration is advantageous in
directing the combustion gases onto the rotatable rotor blades in
the turbine stage 20, it does create additional cooling
difficulties due to the high rate of change of curvature in
transitioning from leading edge 50 to surface 52. The velocity of
the combustion gases at and across the leading edge 50 tends to be
relatively low while the velocity across the suction surface 52 may
become supersonic. Accordingly, there is a significant turbulence
effect as the hot gas stream accelerates from the leading edge to
the suction surface.
Applicants have found that film cooling can be made effective
notwithstanding the broad leading edge configuration and without
adversely affecting performance of the nozzle by forming a
plurality of vent holes 58 in the low Mach number region of the
leading edge 50. While the set of holes 58 may be arranged in
various selected patterns, applicants prefer that the holes 58 are
formed as a radially aligned row of curved or arcuate slots through
the leading edge wall. Applicants have found that an arcuately
shaped or curved vent hole formed with a radius R of about 0.675
inches and an injection angle A of about 16.5 degrees, formed by
the intersection of a line extending from vent hole 58 across a
line tangent to blunt leading edge 50, is not only effective to
establish a cooling or insulative film but provides improved
performance over straight vent holes, in contrast to the
aforementioned NASA report, with a blowing ratio in the order of
1.2. Still further, the arcuately shaped vent holes 58 provide more
effective convective cooling since the effective length of the
holes 58 is longer. It is believed that an injection angle up to 25
degrees can be used with the curved cooling holes and with a
blowing ratio of about 1.2 and still provide effective film
cooling. It may be noted that straight vent holes 60 may be
utilized for film cooling in other areas of the airfoil.
In a preferred embodiment, the cooling air vent holes 58 are formed
as slots having a rectangular cross-section of about 24 mils in
width in the axial or gas stream flow direction and a breadth of 55
mils in the radial direction. Center to center spacing of the slots
or holes 58 is about 0.1 inches in the radial direction so that the
spacing between adjacent slots is about 45 mils. The curved slots
58 exit at an angle of about 16.5 degrees (cooling air injection
angle of 16.5 degrees). The slots 58 are desirably formed using
electric discharge machining (EDM) and a spaced, rectangular, EDM
electrode.
The curved holes 58 provide a significant reduction in cooling air
injection angle which can be reduced below the preferred 16.5
degrees allowing for improved film cooling and coverage by the film
for high blowing ratio (greater than 1.0) applications. More radial
surface of the airfoil is covered by the rectangular slot
configuration of the holes 58 than possible with conventional
circular holes. The injection of the coolant in the low Mach number
region of the airfoil at the leading edge establishes a film of
sufficient quality to effectively cool the entire suction side of
the airfoil. The curved slots 58 provide more effective convective
cooling in the leading edge region of the airfoil.
The degree of curvature in transitioning from the leading edge 50
to the convex suction surface 52 can be appreciated by reference to
the included angle B defined by a line 62 tangent to one of the
arcuate holes 58 and a line 64 tangent to the trailing edge 56. In
the prior art vane airfoils such as that shown in FIG. 3 with the
same tangent lines, the included angle B' is obtuse, typically
being greater than 125 degrees. In the vane of FIG. 4, the included
angle B is acute and typically about 80 degrees.
While other cooling air injection holes, indicated generally at 60,
have not been discussed herein, it will be appreciated that the
airfoil includes such other cooling air holes and that such other
holes may be formed and positioned in a manner similar to the prior
art. The forming and positioning of such other holes 60 is not
significantly different since such other holes are positioned
downstream of the high curvature region and below the blunt leading
edge 50.
What has been disclosed is an improved film cooling method and
apparatus for a blunt leading edge airfoil. While the invention has
been described in what is presently considered to be a preferred
embodiment, various modifications and improvements will become
apparent to those skilled in the art. It is intended therefore that
the invention not be limited to the specific embodiment but be
interpreted within the full spirit and scope of the appended
claims.
* * * * *